Turbine airfoil segment having film cooling hole arrangement

ABSTRACT

A turbine airfoil segment includes inner and outer platforms that are joined by at least one airfoil. The airfoil includes leading and trailing edges that are joined by spaced apart first and second sides to provide an exterior airfoil surface. The airfoil includes film cooling holes that have external breakout points that are located in substantial conformance with the Cartesian coordinates set forth in one of Tables 1 and 2. The Cartesian coordinates are provided by an axial coordinate, a circumferential coordinate, and a radial coordinate, relative to a zero-coordinate, and the film cooling holes have a diametrical surface tolerance relative to the specified coordinates of 0.20 inches (5.0 mm).

CROSS-REFERENCE TO RELATED APPLICATION

The present disclosure claims priority to U.S. Provisional PatentApplication No. 62/088,926, filed Dec. 8, 2014.

BACKGROUND

This disclosure relates to a gas turbine engine and, more particularly,to a turbine airfoil segment that may be incorporated into a gas turbineengine.

Gas turbine engines typically include a compressor section, a combustorsection and a turbine section. During operation, air is pressurized inthe compressor section and is mixed with fuel and burned in thecombustor section to generate hot combustion gases. The hot combustiongases are communicated through the turbine section, which extractsenergy from the hot combustion gases to power the compressor section andother gas turbine engine loads.

Both the compressor and turbine sections may include alternating seriesof rotating blades and stationary vanes that extend into the core flowpath of the gas turbine engine. For example, in the turbine section,turbine blades rotate and extract energy from the hot combustion gasesthat are communicated along the core flow path of the gas turbineengine. The turbine vanes, which generally do not rotate, guide theairflow for the next set of blades. The turbine vanes can be provided inarc segments that each include one or more airfoils that radially extendbetween inner and outer platforms or endwalls. Blades and vanes aregenerally referred to as “airfoils.”

Turbine vanes and blades can include film cooling features to provide aboundary layer of cooling fluid along external surfaces, which protectsthe airfoil from the hot combustion gases in the core flow path.Non-linear flow analyses and complex strain modeling are required toachieve good cooling, making practical results difficult to predict.Loading and temperature considerations also impose substantial designlimitations, which cannot easily be generalized from one system toanother.

SUMMARY

A turbine airfoil segment according to an example of the presentdisclosure includes inner and outer platforms that are joined by atleast one airfoil. The airfoil includes leading and trailing edges thatare joined by spaced apart first and second sides to provide an exteriorairfoil surface, and the at least one airfoil includes film coolingholes that have external breakout points that are located in substantialconformance with the Cartesian coordinates set forth in one of Tables 1and 2. The Cartesian coordinates provided by an axial coordinate, acircumferential coordinate, and a radial coordinate, relative to azero-coordinate, and the film cooling holes have a diametrical surfacetolerance relative to the specified coordinates of 0.20 inches (5.0 mm).

In a further embodiment of any of the foregoing embodiments, the atleast one airfoil includes first and second airfoils, the externalbreakout points of the film cooling holes of the first airfoil arelocated in substantial conformance with the Cartesian coordinates setforth in Table 1 and the external breakout points of the film coolingholes of the second airfoil are located in substantial conformance withthe Cartesian coordinates set forth in Table 2.

In a further embodiment of any of the foregoing embodiments, a portionof the film cooling holes are diffusing holes and another portion of thefilm cooling holes are cylindrical holes.

In a further embodiment of any of the foregoing embodiments, theexternal breakout points of the film cooling holes of the at least oneairfoil are located in substantial conformance with the Cartesiancoordinates set forth in Table 1.

In a further embodiment of any of the foregoing embodiments, the filmcooling holes numbered 28 through 41 in Table 1 are diffusing holes.

In a further embodiment of any of the foregoing embodiments, the filmcooling holes numbered 1 through 27 in Table 1 are cylindrical holes.

In a further embodiment of any of the foregoing embodiments, theexternal breakout points of the film cooling holes of the at least oneairfoil are located in substantial conformance with the Cartesiancoordinates set forth in Table 2.

In a further embodiment of any of the foregoing embodiments, the filmcooling holes numbered 28 through 43 in Table 2 are diffusing holes andthe film cooling holes numbered 1 through 27 in Table 2 are cylindricalholes.

In a further embodiment of any of the foregoing embodiments, spacingbetween edges of adjacent cooling holes is at least 0.015 inch (0.38mm).

In a further embodiment of any of the foregoing embodiments, the filmcooling holes have a diameter of 0.010-0.035 inch (0.25-0.89 mm).

A gas turbine engine according to an example of the present disclosureincludes a compressor section, a combustor fluidly connected to thecompressor section, and a turbine section fluidly connected to thecombustor. The turbine section includes an array of turbine airfoilsegments, each turbine airfoil segment including inner and outerplatforms that are joined by at least one airfoil. The airfoil includesleading and trailing edges that are joined by spaced apart first andsecond sides to provide an exterior airfoil surface, and the at leastone airfoil includes film cooling holes that have external breakoutpoints that are located in substantial conformance with the Cartesiancoordinates set forth in one of Tables 1 and 2. The Cartesiancoordinates provided by an axial coordinate, a circumferentialcoordinate, and a radial coordinate, relative to a zero-coordinate, andthe film cooling holes have a diametrical surface tolerance relative tothe specified coordinates of 0.20 inches (5.0 mm).

In a further embodiment of any of the foregoing embodiments, the atleast one airfoil includes first and second airfoils, the externalbreakout points of the film cooling holes of the first airfoil arelocated in substantial conformance with the Cartesian coordinates setforth in Table 1 and the external breakout points of the film coolingholes of the second airfoil are located in substantial conformance withthe Cartesian coordinates set forth in Table 2.

In a further embodiment of any of the foregoing embodiments, a portionof the film cooling holes are diffusing holes and another portion of thefilm cooling holes are cylindrical holes.

A turbine airfoil segment according to an example of the presentdisclosure includes inner and outer platforms that are joined by firstand second airfoils. Each of the first and second airfoils includesleading and trailing edges that are joined by spaced apart first andsecond sides to provide an exterior airfoil surface. The first andsecond airfoils include, respectively, first and second sets of filmcooling holes that have external breakout points that are located insubstantial conformance with the Cartesian coordinates set forth inTables 1 and 2. The Cartesian coordinates provided by an axialcoordinate, a circumferential coordinate, and a radial coordinate,relative to a zero-coordinate, and the film cooling holes have adiametrical surface tolerance relative to the specified coordinates of0.20 inches (5.0 mm).

In a further embodiment of any of the foregoing embodiments, the filmcooling holes numbered 28 through 41 in Table 1 are diffusing holes.

In a further embodiment of any of the foregoing embodiments, the filmcooling holes numbered 1 through 27 in Table 1 are cylindrical holes.

In a further embodiment of any of the foregoing embodiments, the filmcooling holes numbered 28 through 43 in Table 2 are diffusing holes andthe film cooling holes numbered 1 through 27 in Table 2 are cylindricalholes.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of the present disclosure willbecome apparent to those skilled in the art from the following detaileddescription. The drawings that accompany the detailed description can bebriefly described as follows.

FIG. 1 illustrates an example gas turbine engine.

FIG. 2 illustrates selected portions of a high pressure turbine of thegas turbine engine.

FIG. 3 illustrates a view from the front of a turbine airfoil segment ofthe high pressure turbine.

FIG. 4 illustrates a view from behind the turbine airfoil segment ofFIG. 3.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative enginedesigns can include an augmentor section (not shown) among other systemsor features.

The fan section 22 drives air along a bypass flow path B in a bypassduct defined within a nacelle 15, while the compressor section 24 drivesair along a core flow path C for compression and communication into thecombustor section 26 then expansion through the turbine section 28.Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, the examples herein are not limitedto use with two-spool turbofans and may be applied to other types ofturbomachinery, including direct drive engine architectures, three-spoolengine architectures, and ground-based turbines.

The engine 20 generally includes a low speed spool 30 and a high speedspool 32 mounted for rotation about an engine central longitudinal axisA relative to an engine static structure 36 via several bearing systems38. It should be understood that various bearing systems 38 at variouslocations may alternatively or additionally be provided, and thelocation of bearing systems 38 may be varied as appropriate to theapplication.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48, to drivethe fan 42 at a lower speed than the low speed spool 30.

The high speed spool 32 includes an outer shaft 50 that interconnects asecond (or high) pressure compressor 52 and a second (or high) pressureturbine 54. A combustor 56 is arranged between the high pressurecompressor 52 and the high pressure turbine 54. A mid-turbine frame 57of the engine static structure 36 is arranged generally between the highpressure turbine 54 and the low pressure turbine 46. The mid-turbineframe 57 further supports the bearing systems 38 in the turbine section28. The inner shaft 40 and the outer shaft 50 are concentric and rotatevia bearing systems 38 about the engine central longitudinal axis A,which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines, including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tram °R)/(518.7°R)]^(0.5). The “Low corrected fan tip speed” as disclosed hereinaccording to one non-limiting embodiment is less than about 1150ft/second.

In a further example, the fan 42 includes less than about 26 fan blades.In another non-limiting embodiment, the fan 42 includes less than about20 fan blades. Moreover, in one further embodiment the low pressureturbine 46 includes no more than about 6 turbine rotors schematicallyindicated at 46 a. In a further non-limiting example the low pressureturbine 46 includes about 3 turbine rotors. A ratio between the numberof blades of the fan 42 and the number of low pressure turbine rotors 46a is between about 3.3 and about 8.6. The example low pressure turbine46 provides the driving power to rotate the fan section 22 and thereforethe relationship between the number of turbine rotors 46 a in the lowpressure turbine 46 and the number of blades in the fan section 22discloses an example gas turbine engine 20 with increased power transferefficiency.

FIG. 2 illustrates a cross-sectional view of a portion of the highpressure turbine section 54. The high pressure turbine section 54includes first and second arrays 54 a, 54 c of circumferentially spacedfixed vanes 60, 62. The arrays 54 a, 54 c are axially spaced apart fromone another. A first stage array 54 b of circumferentially spacedturbine blades 64, mounted to a rotor disk 68, is arranged axiallybetween the first and second fixed vane arrays 54 a, 54 c. A secondstage array 54 d of circumferentially spaced turbine blades 66 isarranged aft of the second array 54 c of fixed vanes 62. A platform 58of the second fixed vane array 62 is arranged in an overlappingrelationship with the turbine blades 64, 66.

The turbine blades 64, 66 each include a free tip end 70 adjacent to ablade outer air seal 72 of a case structure 74. The first and secondstage arrays 54 a, 54 c of turbine vanes and first and second stagearrays 54 b, 54 d of turbine blades are arranged within the core flowpath C and are operatively connected to the high speed spool 32. Thesecond stage arrays 54 c of turbine vanes includes a plurality ofturbine airfoil segments 80. Each segment 80 provides an arc length suchthat the segments 80 together provide a complete ring around the enginecentral longitudinal axis A.

FIGS. 3 and 4 show additional, isolated views of a representative one ofthe segments 80. The segment 80 includes inner and outer platforms 82,84 that are joined by at least one airfoil 86. In this example, thesegment 80 is a vane “doublet” and includes two such airfoils 86, namelyfirst airfoil 86 a and second airfoil 86 b. It is to be understood,however, that the segment 80 could alternatively include only a singleairfoil 86, either first airfoil 86 a and second airfoil 86 b, oradditional airfoils 86.

Each airfoil 86 includes leading and trailing edges (represented at “LE”and “TE”) that are joined by spaced apart first and second sides 88 a,88 b to provide an exterior airfoil surface. Sometimes the first andsecond sides 88 a, 88 b are referred to as pressure and suction sides.

Each segment 80 can be formed of a high strength, heat resistantmaterial, such as but not limited to a nickel-based or cobalt-basedsuperalloy, or a high temperature, stress-resistant ceramic or ceramiccomposite material. In cooled configurations, internal fluid passagesand external cooling apertures provide for a combination of convectionand film cooling. In addition, one or more thermal barrier coatings,abrasion-resistant coatings or other protective coatings may be appliedto the segments 80, or at least portions thereof.

Each airfoil 86 also includes film cooling holes, generally representedat 90, that have external breakout points that are located insubstantial conformance with the Cartesian coordinates set forth in oneof Tables 1 and 2 below. The Cartesian coordinates are provided by anaxial coordinate (X-coordinate), a circumferential coordinate(Y-coordinate), and a radial coordinate (Z-coordinate), relative to azero-coordinate. The axial coordinate is along a direction parallel tothe engine axis A. The radial coordinate is along a directionperpendicular to the engine axis A, and the circumferential coordinateis along a circumferential direction about the engine axis A. In oneexample, the zero-coordinate is at point “Pt” located with respect tothe curvature of the arc of the segment 80. In one example, the point“Pt” is located at the centerpoint of the curvature of arc surface ASand on a plane coincident with surface P.

The coordinates of Tables 1 and 2 (in inches) provide the nominal axial,circumferential, and radial coordinates relative to the zero-coordinate,on a cold, uncoated, stationary segment 80. Each row in Tables 1 and 2corresponds to a single film cooling hole 90 location. Additionalelements, such as additional cooling holes, protective coatings, filletsand seal structures may also be formed onto the external surfaces of theairfoils 86, but these elements are not necessarily described by thecoordinates.

Due to manufacturing tolerances, the film cooling holes 90 have adiametrical surface tolerance, relative to the specified coordinates, of0.20 inches (5.0 mm). That is, there is a spatial envelope in which thefilm cooling hole 90 is located. In a further example, a minimum spacingis provided between adjacent film cooling holes 90. In one example, theminimum spacing between edges of adjacent film cooling holes 90 is atleast 0.015 inch (0.38 mm).

The film cooling holes 90 are arranged to produce film of cooling fluidon the external surfaces of the airfoils 86. As shown, portions of thefilm cooling holes 90 are arranged in clusters or radial rows to providefilm cooling at particular locations. In Tables 1 and 2, each filmcooling hole has a Row ID and a hole number. The Row ID nomenclature hasthree letters. The first two letter designate a row and the last letterdesignates the hole of that row (e.g., holes A through N in cluster HA).Table 2 uses a similar nomenclature.

The film cooling holes 90, or clusters of holes, can be diffusing holesor cylindrical holes, for example, but are not limited to suchgeometries. In diffusing hole geometries, the hole diameter areaincreases as the hole opens to the external surface. Cylindrical holeshave a uniform diameter area along the length of the hole. In furtherexamples, the film cooling holes numbered 28 through 41 in Table 1 arediffusing holes and the film cooling holes numbered 1 through 27 inTable 1 are cylindrical holes. In a further example, the film coolingholes numbered 28 through 43 in Table 2 are diffusing holes and the filmcooling holes numbered 1 through 27 in Table 2 are cylindrical holes.

Diffusing holes can provide good film coverage in comparison with acylindrical hole of the same size. Diffusing holes can be used whereenhanced cooling is desired. Cylindrical holes can provide highervelocity cooling flow in comparison to conical holes of the same size.In one further example, the film cooling holes 90 have a minimumdiameter of 0.010-0.035 inch (0.25-0.89 mm).

TABLE 1 First Airfoil Row ID Hole ID X Y Z HAA 1 −1.778 −0.525 8.076 HAB2 −1.777 −0.538 8.182 HAC 3 −1.778 −0.551 8.288 HAD 4 −1.779 −0.5618.395 HAE 5 −1.781 −0.569 8.500 HAF 6 −1.783 −0.576 8.606 HAG 7 −1.783−0.579 8.712 HAH 8 −1.785 −0.580 8.818 HAJ 9 −1.784 −0.576 8.924 HAK 10−1.784 −0.571 9.030 HAL 11 −1.784 −0.565 9.136 HAM 12 −1.783 −0.5599.238 HAN 13 −1.783 −0.553 9.329 HBA 14 −1.781 −0.631 8.124 HBB 15−1.779 −0.644 8.231 HBC 16 −1.776 −0.656 8.338 HBD 17 −1.773 −0.6658.453 HBE 18 −1.771 −0.672 8.551 HBF 19 −1.769 −0.676 8.658 HBG 20−1.768 −0.677 8.765 HBH 21 −1.768 −0.675 8.873 HBJ 22 −1.769 −0.6728.978 HBK 23 −1.770 −0.667 9.086 HBL 24 −1.770 −0.661 9.190 HBM 25−1.771 −0.655 9.287 HBN 26 −1.772 −0.649 9.381 HCA 27 −1.793 −0.6208.476 PBA 28 −0.772 0.667 7.993 PBB 29 −0.769 0.672 8.101 PBC 30 −0.7630.676 8.208 PBD 31 −0.755 0.677 8.316 PBE 32 −0.745 0.681 8.423 PBF 33−0.737 0.682 8.530 PBG 34 −0.728 0.684 8.636 PBH 35 −0.720 0.689 8.743PBJ 36 −0.711 0.699 8.850 PBK 37 −0.704 0.708 8.957 PBL 38 −0.699 0.7189.064 PBM 39 −0.695 0.727 9.170 PBN 40 −0.690 0.738 9.277 PBP 41 −0.6860.749 9.384

TABLE 2 Second Airfoil Row ID Hole ID X Y Z HAA 1 −1.778 −0.525 8.076HAB 2 −1.777 −0.538 8.182 HAC 3 −1.778 −0.551 8.288 HAD 4 −1.779 −0.5618.395 HAE 5 −1.781 −0.569 8.500 HAF 6 −1.783 −0.576 8.606 HAG 7 −1.783−0.579 8.712 HAH 8 −1.785 −0.580 8.818 HAJ 9 −1.784 −0.576 8.924 HAK 10−1.784 −0.571 9.030 HAL 11 −1.784 −0.565 9.136 HAM 12 −1.783 −0.5599.238 HAN 13 −1.783 −0.553 9.329 HBA 14 −1.781 −0.631 8.124 HBB 15−1.779 −0.644 8.231 HBC 16 −1.776 −0.656 8.338 HBD 17 −1.773 −0.6658.453 HBE 18 −1.771 −0.672 8.551 HBF 19 −1.769 −0.676 8.658 HBG 20−1.768 −0.677 8.765 HBH 21 −1.768 −0.675 8.873 HBJ 22 −1.769 −0.6728.978 HBK 23 −1.770 −0.667 9.086 HBL 24 −1.770 −0.661 9.190 HBM 25−1.771 −0.655 9.287 HBN 26 −1.772 −0.649 9.381 HCA 27 −1.793 −0.6208.476 PAA 28 −0.776 0.661 7.944 PAB 29 −0.767 0.671 8.157 PAC 30 −0.7500.680 8.370 PAD 31 −0.732 0.684 8.582 PAE 32 −0.717 0.691 8.795 PAF 33−0.706 0.705 9.007 PAG 34 −0.697 0.723 9.219 PAH 35 −0.683 0.755 9.429SAA 36 −1.101 −0.500 8.194 SAB 37 −1.104 −0.519 8.342 SAC 38 −1.106−0.534 8.489 SAD 39 −1.102 −0.540 8.637 SAE 40 −1.102 −0.547 8.786 SAF41 −1.102 −0.552 8.934 SAG 42 −1.105 −0.561 9.082 SAH 43 −1.110 −0.5729.230

Substantial conformance with the coordinates of Tables 1 and 2 is basedon points representing the film cooling hole 90 locations, for examplein inches or millimeters, as determined by selecting particular valuesof scaling parameters. A substantially conforming segment has filmcooling holes that conform to the specified sets of points, within thespecified tolerance.

Alternatively, substantial conformance is based on a determination by anational or international regulatory body, for example in a partcertification or part manufacture approval (PMA) process for the FederalAviation Administration, the European Aviation Safety Agency, the CivilAviation Administration of China, the Japan Civil Aviation Bureau, orthe Russian Federal Agency for Air Transport. In these configurations,substantial conformance encompasses a determination that a particularpart or structure is identical to, or sufficiently similar to, thespecified vane, or that the part or structure is sufficiently the samewith respect to a part design in a type-certified or type-certificatedvane, such that the part or structure complies with airworthinessstandards applicable to the specified vane. In particular, substantialconformance encompasses any regulatory determination that a particularpart or structure is sufficiently similar to, identical to, or the sameas a specified vane, such that certification or authorization for use isbased at least in part on the determination of similarity.

Although a combination of features is shown in the illustrated examples,not all of them need to be combined to realize the benefits of variousembodiments of this disclosure. In other words, a system designedaccording to an embodiment of this disclosure will not necessarilyinclude all of the features shown in any one of the Figures or all ofthe portions schematically shown in the Figures. Moreover, selectedfeatures of one example embodiment may be combined with selectedfeatures of other example embodiments.

The preceding description is exemplary rather than limiting in nature.Variations and modifications to the disclosed examples may becomeapparent to those skilled in the art that do not necessarily depart fromthe essence of this disclosure. The scope of legal protection given tothis disclosure can only be determined by studying the following claims.

What is claimed is:
 1. A turbine airfoil segment comprising: inner andouter platforms that are joined by at least one airfoil to form asegment that has an arc, the at least one airfoil includes leading andtrailing edges that are joined by spaced apart first and second sides toprovide an exterior airfoil surface, and the at least one airfoilincludes film cooling holes that have external breakout points that arelocated in Cartesian coordinates set forth in one of Tables 1 and 2, theCartesian coordinates provided by an axial coordinate, a circumferentialcoordinate, and a radial coordinate, relative to a zero-coordinate,wherein the zero-coordinate is at point, Pt, located with respect to acurvature of the arc of the segment, and the film cooling holes have adiametrical surface tolerance to the specified coordinates of 0.20inches (5.0 mm).
 2. The turbine airfoil segment as recited in claim 1,wherein the at least one airfoil includes first and second airfoils, theexternal breakout points of the film cooling holes of the first airfoilare located in the Cartesian coordinates set forth in Table 1 and theexternal breakout points of the film cooling holes of the second airfoilare located in the Cartesian coordinates set forth in Table
 2. 3. Theturbine airfoil segment as recited in claim 2, wherein a portion of thefilm cooling holes are diffusing holes and another portion of the filmcooling holes are cylindrical holes.
 4. The turbine airfoil segment asrecited in claim 1, wherein the external breakout points of the filmcooling holes of the at least one airfoil are located in the Cartesiancoordinates set forth in Table
 1. 5. The turbine airfoil segment asrecited in claim 4, wherein the film cooling holes numbered 28 through41 in Table 1 are diffusing holes.
 6. The turbine airfoil segment asrecited in claim 5, wherein the film cooling holes numbered 1 through 27in Table 1 are cylindrical holes.
 7. The turbine airfoil segment asrecited in claim 1, wherein the external breakout points of the filmcooling holes of the at least one airfoil are located in the Cartesiancoordinates set forth in Table
 2. 8. The turbine airfoil segment asrecited in claim 7, wherein the film cooling holes numbered 28 through43 in Table 2 are diffusing holes and the film cooling holes numbered 1through 27 in Table 2 are cylindrical holes.
 9. The turbine airfoilsegment as recited in claim 1, wherein the film cooling holes have adiameter of 0.010-0.035 inch (0.25-0.89 mm).
 10. A gas turbine enginecomprising: a compressor section; a combustor fluidly connected to thecompressor section; a turbine section fluidly connected to thecombustor, the turbine section includes an array of turbine airfoilsegments, each turbine airfoil segment having an arc and comprising:inner and outer platforms that are joined by at least one airfoil, theat least one airfoil includes leading and trailing edges that are joinedby spaced apart first and second sides to provide an exterior airfoilsurface, and the at least one airfoil includes film cooling holes thathave external breakout points that are located in Cartesian coordinatesset forth in one of Tables 1 and 2, the Cartesian coordinates providedby an axial coordinate, a circumferential coordinate, and a radialcoordinate, relative to a zero-coordinate, wherein the zero-coordinateis at point, Pt, located with respect to a curvature of the arc of theturbine airfoil segment, and the film cooling holes have a diametricalsurface tolerance to the specified coordinates of 0.20 inches (5.0 mm).11. The gas turbine engine as recited in claim 10, wherein the at leastone airfoil includes first and second airfoils, the external breakoutpoints of the film cooling holes of the first airfoil are located in theCartesian coordinates set forth in Table 1 and the external breakoutpoints of the film cooling holes of the second airfoil are located inthe Cartesian coordinates set forth in Table
 2. 12. The gas turbineengine as recited in claim 11, wherein a portion of the film coolingholes are diffusing holes and another portion of the film cooling holesare cylindrical holes.
 13. A turbine airfoil segment comprising: innerand outer platforms that are joined by first and second airfoils to forma segment that has an arc, each of the first and second airfoilsincludes leading and trailing edges that are joined by spaced apartfirst and second sides to provide an exterior airfoil surface, and thefirst and second airfoils include, respectively, first and second setsof film cooling holes that have external breakout points that arelocated in Cartesian coordinates set forth in Tables 1 and 2, theCartesian coordinates provided by an axial coordinate, a circumferentialcoordinate, and a radial coordinate, relative to a zero-coordinate,wherein the zero-coordinate is at point, Pt, located with respect to acurvature of the arc of the segment, and the film cooling holes have adiametrical surface tolerance to the specified coordinates of 0.20inches (5.0 mm).
 14. The turbine airfoil segment as recited in claim 13,wherein the film cooling holes numbered 28 through 41 in Table 1 arediffusing holes.
 15. The turbine airfoil segment as recited in claim 14,wherein the film cooling holes numbered 1 through 27 in Table 1 arecylindrical holes.
 16. The turbine airfoil segment as recited in claim15, wherein the film cooling holes numbered 28 through 43 in Table 2 arediffusing holes and the film cooling holes numbered 1 through 27 inTable 2 are cylindrical holes.